In a gas turbine engine, ambient air is drawn into a compressor section. Alternate rows of stationary and rotating aerofoil blades are arranged around a common axis; together these accelerate and compress the incoming air. A rotating shaft drives the rotating blades. Compressed air is delivered to a combustor section where it is mixed with fuel and ignited. Ignition causes rapid expansion of the fuel/air mix which is directed in part to propel a body carrying the engine and in another part to drive rotation of a series of turbines arranged downstream of the combustor. The turbines share rotor shafts in common with the rotating blades of the compressor and work, through the shaft, to drive rotation of the compressor blades.
It is well known that the operating efficiency of a gas turbine engine is improved by increasing the operating temperature. The ability to optimise efficiency through increased temperatures is restricted by changes in behaviour of materials used in the engine components at elevated temperatures which, amongst other things, can impact upon the mechanical strength of the blades and the rotor disc which carries the blades. This problem is addressed by providing a flow of coolant through and/or over the turbine rotor disc and blades.
It is known to take off a portion of the air output from the compressor (which is not subjected to ignition in the combustor and so is relatively cooler) and feed this to surfaces in the turbine section which are likely to suffer damage from excessive heat. Typically the cooling air is delivered adjacent the rim of the turbine disc and directed to a port which enters the turbine blade body and is distributed through the blade, typically by means of a labyrinth of channels extending through the blade body. Heat transfers between this cooling air flow and the turbine disc can increase thermal gradients in the disc leading to the disc material being subjected to an increased stress range. This stress range is a limiting factor in the life of the disc.
Various component bucket groove heat shields are known. An example is described in the Applicant's prior published U.S. Pat. No. 8,845,288B. Such arrangements involve complex and time consuming assembly and often suffer from leakage of coolant air into the bucket groove.
It is known to provide turbine blades which incorporate an integral heat shield at the root. The blade root is shaped to fit the contours of a bucket groove which extends radially inwardly of the disc rim. A duct (into which cooling air for the blade is to be received) is machined into the blade root leaving a duct wall adjacent the radially innermost surface of the bucket groove. The duct wall serves as a heat shield between the disc body and the coolant flow. An example is described in Applicant's co-pending European patent application no. EP16162567.8. Whilst effective, such heat shields present complex manufacturing challenges. Cooling channels extending from the duct through the blade body have carefully designed inlet geometries and dimensions which serve to encourage dominant flow to regions of the blade which most need it. In addition it may be desirable to apply finishing operations such as coating to the channels. The presence of the integral shield presents an obstacle to the performance of such finishing operations.
An objective of the present disclosure is to mitigate issues with the prior art arrangements.